LEE- Uçak ve Uzay Mühendisliği-Yüksek Lisans

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http://hdl.handle.net/11527/19474

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A metamodel based approach for the shape optimization of a front rail

(Graduate School, 2024-12-23) Gülmüş, Mustafa ; Türkmen, Halit Süleyman ; 511211177 ; Aeronautics and Astronautics Engineering

In crash cases, injuries and loss of lifes are inevitable situation most of the time. There are some another considerations in economic, sustainability and production areas from the manufacturer's point of view. There have been some technologic developments in automobile and aerospace industry that pushes manufacturers to produce safer and more profitable products. Bumper, rocker, rail upper and front rail are the main crashworthiness structures for vehicles design. Front rail is one of the most important impact absorber among the all crashworthiness structures in vehicle design area. It is noted that front rail plays the most critical role during the crash in crashworthiness structures. The fuctions of front rail are transmissing the crash force to the middle crumple zone, absorbing impact energy and providing a favorable buckling form during the crash. Absorbing impact energy is the most important parameter among all functions for passenger safety. Before the crash, there is a high kinetic energy for passenger between vehicles or between vehicle and static objects. One of the most considerable condition for crash cases is passenger safety. Crashworthiness structures are the main provider for this case. They absorbs the some portion of kinetic energy and transmits the left part into the passenger zone. It is important for passenger safety that the more energy absorbed by crashworthiness structures because passengers are exposed less impact energy. There have been some optimization studies and design innovations for crashworthiness structures. Also, some material applications are made to overcome in absorbing impact energy consideration for crashworthiness structures. Addition of sub-structures into crashworthiness structures, optimizations for getting less weight structures, structural optimizations are some studies for crashworthiness structures. In real crash test centers, speed of vehicles are specified as encountered in urban areas. Most of the crashes happens in city streets and real tests center used these kind of speed levels. Design modifications, material applications and comparing cases for different speed levels are the main objectives in this study. Firstly, a design approximation is made to absorb more impact energy in optimization cases. Embosses are created on the front rail for this purpose. Relations between embosses and thickness of front rail are specified as design variables. Another purpose of creating embosses on the front rail is getting appropriate buckling form during the crash. Embosses are provides this problem in optimization study. Secondly, some material applications are made for economic and flammability considerations. In crash scenarios, a fire situation can come up and it might melt the crashworthiness structures. Another consideration is fuel consumption. Selecting different materials provide us to have lightweight structure. Specific Energy Absorption (SEA) is another parameter for material selection for crash worthiness structure. Three different materials are used for every optimization process in this study. Lastly, two different speeds are selected to see how front rail behaves under low and high speed conditions and their buckling forms are examined. A stochastic optimizaiton method is selected for optimization processes. Genetic algorithm is used for all optimization studies. Metamodel optimization technics are used with the usage of genetic algorithm. D-optimal point selection is used for sampling process and sequantial response with domain reduction method is used for the selection of next iteration sampling design space. In result section, all six cases are compared with each other for crashworthiness indicators like SEA, peak force, mean crash force et al. After all optimization processes importance of design variables are discussed and metamodels of optimization processes are shown. Von-mises stress distributions of front rails are shown for all six cases to visualize stress distribution. Advantages and disadvantage of all six cases are discussed and their efficiency on this study are indicated.
A multiscale approach to understand the effects of design parameters on the elastic behavior of 3D orthogonally woven composites

(Graduate School, 2024-11-04) Erkoç, Hilal ; Cebeci, Hülya ; 511201167 ; Aeronautical and Astronautical Engineering

This study aims to investigate the effect of various parameters on the elastic constants of three-dimensional (3D) orthogonally woven composites. Two-dimensional (2D) laminated composites exhibit high in-plane stiffness and strength; however, they are inadequate in applications subjected to out-of-plane loads, particularly in engine fan blades, aircraft fuselage structures, and wind turbine blades. With an innovative approach, 3D orthogonally woven composites effectively overcome the limitations of traditional 2D laminates. The usage of 3D orthogonally woven composites in these structures can be beneficial because 3D orthogonally woven composites are more resistant to out-of-plane loading than 2D laminates, due to their improved mechanical properties through the thickness. In addition to this, improved impact damage tolerance, higher delamination resistance, and reduced assembly and production costs through single-piece fabric production are advantages of 3D orthogonally woven composites. 3D orthogonally woven composites, in spite of their advantages, present certain challenges in application. One of the significant challenges is the complex nature of their manufacturing process, which demands specialized equipment and skilled personnel, leading to high production costs. Their complex structure can also complicate design, analysis, and simulation, requiring advanced computational models. Additionally, the complex architecture of these composites can present challenges in repair and maintenance procedures. 3D orthogonally woven structures consist of three interwoven sets of yarns arranged in orthogonal directions, where the warp and weft yarns remain straight while the binder yarns interlace them to create a multidimensional architecture. This complex architecture of 3D orthogonally woven composites plays an important role in determining the mechanical properties of the structure. Since differences in cross-section configurations, yarn arrangements, and fiber interactions significantly influence the load-carrying capacity, stiffness, and overall performance of the composite, an in-depth examination of the structural architecture is critical to optimizing the mechanical properties of the material. Several analytical studies have examined the effects of binder-to-weft and binder-to-warp ratios on the elastic properties of 3D orthogonally woven composites. These analyses employ representative volume elements (RVEs) to model the material behaviors. The binder-to-weft ratio characterizes the number of wefts of yarn a binder yarn encircles before reversing direction within the weft layer. Similarly, the binder-to-warp ratio represents the proportion of warp yarns per layer relative to the total number of warps encompassed by the RVE. However, a key limitation of these existing studies is based on the absence of a comparative analysis between analytical solutions and numerical simulations. Furthermore, the impact of RVE thickness on its elastic coefficients has not been thoroughly investigated. Here, the effects of changing thickness on the tensile response of the structure, as obtained through analytical solutions and numerical simulations, are presented. Elastic constants of 3D fiber-reinforced composites were estimated using a multi-scale homogenization technique based on meso-macro homogenization with good correlation. Numerical simulations were performed using ABAQUS software to analyze the behavior of the models. Through the optimization of the geometrical parameters of RVE, 3D orthogonally woven composites can be effectively implemented across a diverse range of engineering applications, especially in the aviation field.
A numerical investigation of total temperature probes measurement performance

(Graduate School, 2023) Meriç, Erdem ; Edis, Fırat Oğuz ; 810043 ; Aeronautics and Astronautics Engineering Program

In almost every industrial application, the temperature is measured for development and condition monitoring purposes. The accuracy of these measurements is crucial to avoid misunderstandings about the current condition and misguidance in the development phase. The most practical mean of temperature measurement in industrial applications is using a thermocouple. Thermocouples are very flexible structures so they can be applied in many different regions for solid and fluid temperature measurements. It is also possible to design measurement probe geometries using thermocouples as sensing elements. In machines involving high-speed gas flow, the kinetic energy of fluid can't be neglected in energy interaction calculations so flow must be adiabatically stagnated before temperature measurement. The temperature a flowing fluid gains because of adiabatic stagnation is called stagnation or total temperature. A stationary probe geometry measures the total temperature of flow but there may be deviations in the temperature of the sensing point due to the flow physics. These deviations lead to errors in measurement. These errors are classified as recovery error, conduction error and radiation error. Recovery error originated from the non-adiabatic stagnation of flow on the surface of the thermocouple (TC) junction. Recovery error is characterized by a parameter called recovery factor which shows the degree of dynamic temperature recovery on the measurement. Conduction and radiation errors arise due to solid boundary conditions which are different from the flow total temperature around the probe. These different temperature zones cause heat interaction via conduction and radiation heat transfer modes between the TC junction and surroundings giving rise to deviations in measurement. Special probe designs are used to prevent these errors. In this study, an experimental case was selected from the literature to create a conjugate heat transfer (CHT) methodology. This CHT methodology served to investigate flow physics around and inside total temperature probes and the nature of heat interaction between flow and probe geometry. This experimental case contains a total temperature probe calibration setup which investigates the measurement performance of probe geometry under different Mach number flows. In the simulations, the measurement probe geometry was modelled and exposed to the flow at the same speed as the test conditions. The main observed parameter during simulations was TC junction temperature which determines the performance of the total temperature probe. The results of simulations were observed to be in harmony with experimental data. Then, flow structures around and inside the total temperature probe were investigated in detail using the outputs of simulations. The main aim of total temperature probe geometry is to decrease flow velocity inside the shield to decrease thermal conduction in the boundary layer. In simulations, this aim was observed to be accomplished. The flow velocity vectors were investigated to understand the nature of flow around and inside the total temperature probe. No flow separation was observed on the shield inlet.
A study on optimization of a wing with fuel sloshing effects

(Graduate School, 2022-01-24) Vergün, Tolga ; Doğan, Vedat Ziya ; 511181206 ; Aeronautics and Astronautics Engineering ; Uçak ve Uzay Mühendisliği

In general, sloshing is defined as a phenomenon that corresponds to the free surface elevation in multiphase flows. It is a movement of liquid inside another object. Sloshing has been studied for centuries. The earliest work [48] was carried out in the literature by Euler in 1761 [17]. Lamb [32] theoretically examined sloshing in 1879. Especially with the development of technology, it has become more important. It appears in many different fields such as aviation, automotive, naval, etc. In the aviation industry, it is considered in fuel tanks. Since outcomes of sloshing may cause instability or damage to the structure, it is one of the concerns about aircraft design. To prevent its adverse effect, one of the most popular solutions is adding baffles into the fuel tank. Still, this solution also comes with a disadvantage: an increase in weight. To minimize the effects of added weight, designers optimize the structure by changing its shape, thickness, material, etc. In this study, a NACA 4412 airfoil-shaped composite wing is used and optimized in terms of safety factor and weight. To do so, an initial composite layup is determined from current designs and advice from literature. When the design of the initial system is completed, the system is imported into a transient solver in the Ansys Workbench environment to perform numerical analysis on the time domain. To achieve more realistic cases, the wing with different fuel tank fill levels (25%, 50%, and 75%) is exposed to aerodynamic loads while the aircraft is rolling, yawing, and dutch rolling. The aircraft is assumed to fly with a constant speed of 60 m/s (~120 knots) to apply aerodynamic loads. Resultant force for 60 m/s airspeed is applied onto the wing surface by 1-Way Fluid-Structure Interaction (1-Way FSI) as a distributed pressure. Using this method, only fluid loads are transferred to the structural system, and the effect of wing deformation on the fluid flow field is neglected. Once gravity effects and aerodynamic loads are applied to the wing structure, displacement is defined as the wing is moving 20 deg/s for 3 seconds for all types of movements. On the other hand, fluid properties are described in the Ansys Fluent environment. Fluent defines the fuel level, fluid properties, computational fluid dynamics (CFD) solver, etc. Once both structural and fluid systems are ready, system coupling can perform 2-Way Fluid-Structure Interaction (2-Way FSI). Using this method, fluid loads and structural deformations are transferred simultaneously at each step. In this method, the structural system transfers displacement to the fluid system while the fluid system transfers pressure to the structural system. After nine analyses, the critical case is determined regarding the safety factor. Critical case, in which system has the lowest minimum safety factor, is found as 75% filled fuel tank while aircraft dutch rolling. After the determination of the critical case, the optimization process is started. During the optimization process, 1-Way FSI is used since the computational cost of the 2-Way FSI method is approximately 35 times that of 1-Way FSI. However, taking less time should not be enough to accept 1-Way FSI as a solution method; the deviation of two methods with each other is also investigated. After this investigation, it was found that the variation between the two methods is about 1% in terms of safety factors for our problem. In the light of this information, 1-Way FSI is preferred to apply both sloshing and aerodynamic loads onto the structure to reduce computational time. After method selection, thickness optimization is started. Ansys Workbench creates a design of experiments (DOE) to examine response surface points. Latin Hypercube Sampling Design (LHSD) is preferred as a DOE method since it generates non-collapsing and space-filling points to create a better response surface. After creating the initial response surface using Genetic Aggregation, the optimization process is started using the Multi-Objective Genetic Algorithm (MOGA). Then, optimum values are verified by analyzing the optimum results in Ansys Workbench. When the optimum results are verified, it is realized that there is a notable deviation in results between optimized and verified results. To minimize the variation, refinement points are added to the response surface. This process is kept going until variation comes under 1%. After finding the optimum results, it is noticed that its precision is too high to maintain manufacturability so that it is rounded into 1% of a millimeter. In the end, final thickness values are verified. As a result, optimum values are found. It is found that weight is decreased from 100.64 kg to 94.35 kg, which means a 6.3% gain in terms of weight, while the minimum safety factor of the system is only reduced from 1.56 to 1.54. At the end of the study, it is concluded that a 6.3% reduction in weight would reflect energy saving.
A study on static and dynamic buckling analysis of thin walled composite cylindrical shells

(Graduate School, 2022-01-24) Özgen, Cansu ; Doğan, Vedat Ziya ; 511171148 ; Aeronautics and Astronautics Engineering ; Uçak ve Uzay Mühendisliği

Thin-walled structures have many useage in many industries. Examples of these fields include: aircraft, spacecraft and rockets can be given. The reason for the use of thin-walled structures is that they have a high strength weight ratio. In order to define a cylinder as thin-walled, the ratio of radius to thickness must be more than 20, and one of the problems encountered in the use of such structures is the problem of buckling. It is possible to define the buckling as a state of instability in the structure under compressive loads. This state of instability can be seen in the load displacement graph as the curve follows two different paths. The possible behaviors; snap through or bifurcation behavior. Compressive loading that cause buckling; there may be an axial load, torsional load, bending load, external pressure. In addition to these loads, buckling may occur due to temperature change. Within the scope of this thesis, the buckling behavior of thin-walled cylinders under axial compression was examined. The cylinder under the axial load indicates some displacement. When the amount of load applied reaches critical level, the structure moves from one state of equilibrium to another. After some point, the structure shows high displacement behavior and loses stiffness. The amount of load that the structure will carry decreases considerably, but the structure continues to carry loads. The behavior of the structure after this point is called post-buckling behavior. The critical load level for the structure can be determined by using finite elements method. Linear eigenvalue analysis can be performed to determine the static buckling load. However, it should be noted here that eigenvalue-eigenvector analysis can only be used to make an approximate estimate of the buckling load and input the resulting buckling shape into nonlinear analyses as a form of imperfection. In addition, it can be preferred to change parameters and compare them, since they are cheaper than other types of analysis. Since the buckling load is highly affected by the imperfection, nonlinear methods with geometric imperfection should be used to estimate a more precise buckling load. It is not possible to identify geometric imperfection in linear eigenvalue analysis. Therefore, a different type of analysis should be selected in order to add imperfection. For example, an analysis model which includes imperfection can be established with the Riks method as a nonlinear static analysis type. Unlike the Newton-Rapson method, the Riks method is capable of backtracking in curves. Thus, it is suitable for use in buckling analysis. In Riks analysis, it is recommended to add imperfection in contrast to linear eigenvalue analysis. Because if the imperfection is added, the problem will be bifurcation problem instead of limit load problem and sharp turns in the graph can cause divergence in analysis. Another nonlinear method of static phenomena is called quasi-static analysis which is used dynamic solver. The important thing to note here is that the inertial effects should be too small to be neglected in the analysis. For this purpose, kinetic energy and internal energy should be compared at the end of the analysis and kinetic energy should be ensured to be negligible levels besides internal energy. Also, if the event is solved in the actual time length, this analysis will be quite expensive. Therefore, the time must be scaled. In order to scale the time correctly, frequency analysis can be performed first and the analysis time can be determined longer than the period corresponding to the first natural frequency. For three analysis methods mentioned within this study, validation studies were carried out with the examples in the literature. As a result of each type of analysis giving consistent results, the effect of parameters on static buckling load was examined, while linear eigenvalue analysis method was used because it was also sufficient for cheaper analysis method and comparison studies. While displacement-controlled analyses were carried out in the static buckling analyses mentioned, load-controlled analyses were performed in the analyses for the determination of dynamic buckling force. As a result of these analyses, they were evaluated according to different dynamic buckling criteria. There are some of the dynamic buckling criteria; Volmir criterion, Budiansky-Roth criterion, Hoff-Bruce criterion, etc. When Budiansky-Roth criterion is used, the first estimated buckling load is applied to the structure and displacement - time graph is drawn. If a major change in displacement is observed, it can be assumed that the structure is dynamically buckled. For Hoff-Bruce criterion, the speed - displacement graph should be drawn. If this graph is not focused in a single area and is drawn in a scattered way, it is considered that the structure has moved to the unstable area. As in static buckling analyses, dynamic buckling analyses were primarily validated with a sample study in the literature. After the analysis methods, the numerical studies were carried out on the effect of some parameters on the buckling load. First, the effect of the stacking sequence of composite layers on the buckling load was examined. In this context, a comprehensive study was carried out, both from which layer has the greatest effect of changing the angle and which angle has the highest buckling load. In addition, the some angle combinations are obtained in accordance with the angle stacking rules found in the literature. For those stacking sequences, buckling forces are calculated with both finite element analyses and analytically. In addition, comparisons were made with different materials. Here, the buckling load is calculated both for cylinders with different masses of the same thickness and for cylinders with different thicknesses with the same mass. Here, the highest force value for cylinders with the same mass is obtained for a uniform composite. In addition, although the highest buckling force was obtained for steel material in the analysis of cylinders of the same thickness, when we look at the ratio of buckling load to mass, the highest value was obtained for composite material. In addition, the ratio of length to diameter and the effect of thickness were also examined. Here, as the length to diameter ratio increases, the buckling load decreases. As the thickness increases, the buckling load increases with the square of the thickness. In addition to the effect of the length to diameter ratio and the effect of thickness, the loading time and the shape of the loading profile are also known in dynamic buckling analysis. In addition, the critical buckling force is affected by imperfections in the structure, which usually occur during the production of the structure. How sensitive the structures are to the imperfection may vary depending on the different parameters. The imperfection can be divided into three different groups as geometric, material and loading. Cylinders under axial load are particularly affected by geometric imperfection. The geometric imperfection can be defined as how far the structure is from a perfect cylindrical structure. It is possible to determine the specified amount of deviation by different measurement methods. Although it is not possible to measure the amount of imperfection for all structures, an idea can be gained about how much imperfection is expected from the studies found in the literature. Both the change in the buckling load on the measured cylinders and the imperfection effect of the buckling load can be measured by adding the measured amount of imperfection to the buckling load calculations. In cases where the amount of imperfection cannot be measured, the finite element can be included in the analysis model as an eigenvector imperfection obtained from linear buckling analysis and the critical buckling load can be calculated for the imperfect structure using nonlinear analysis methods. In this study, studies were carried out on how imperfection sensitivity changes under both static and dynamic loading with different parameters. These parameters are the the length-to-diameter ratio, the effect of the stacking sequence of the composite layers and the added imperfection shape. The most important result obtained in the study on imperfection sensitivity is that the effect of the imperfection on the buckling load is quite high. Even geometric imperfection equal to thickness can cause the buckling load to drop by up to half.
Aeroacoustic investigations for a refrigerator air duct and flow systems

(Graduate School, 2022-02-16) Demir, Hazal Berfin ; Çelik, Bayram ; 511181186 ; Aeronautics and Astronautics Engineering

Noise has become an important public health problem with industrialization, and has become a crucial design problem for engineering. For this reason, noise reduction studies have became the focus, especially in the white goods, automotive and aviation sectors, which requires interaction with human. Among the vehicles and products in the aforementioned sectors, the refrigerators, unlike the others, are located in the center of the living area and work throughout the day. Therefore, possible sound problems are observed more quickly by the users and are found to be disturbing. At this point, the investigation and reduction of the acoustic propagation of existing products by various numerical and experimental methods is a valuable contribution to both industry and literature. Within the scope of this thesis, the freezer compartment of a refrigerator with a No frost cooling system was investigated from an aeroacoustic perspective. The freezer compartment consists of three drawers where food will be placed, an axial fan that provides air flow, an evaporator cover that separates the evaporator pipes and the interior volume, and plastic walls surrounding them. The main source of air flow noise in the system is the axial fan. For this reason, in the first step of the study, solo aeroacoustic examination of the axial fan was made. Afterwards, the entire freezer volume was examined and the study was completed with three different model proposals in which acoustic emission was reduced. The flow field analysis of the axial fan with an operational speed of 1200 rpm was carried out with commercial software ANSYS Fluent. In this numerical model, Shear Stress Transport 𝑘 – 𝜔 turbulence model was used. Governing equations was solved under three-dimensional, transient, viscous, incompressible flow assumptions. The rotation of the fan was defined by the sliding mesh method. The numerical flow solution was validated with experimental volumetric flow rate data. According to the numerical and experimental results, the flow rate of the axial fan under the specified conditions was determined as 19 L/s. A hybrid aeroacoustic model is created by giving the pressure outputs of the flow solution as input to the acoustic model. For the acoustic solution, Ffowcs Williams & Hawkings (FW-H) model defined in ANSYS Fluent was used and the result of the solution was compared with the sound pressure data collected in the full anechoic acoustic room. Although there is some difference between the numerical and experimental sound pressure curves, it was observed that the hybrid model established to understand the general trend and to catch the blade passing frequency was successful. It was predicted that the difference between experimental and numerical measurements occurred for two reasons. The first is absence of the fan motor in the numerical analysis. Another reason is that the acoustic propagation resulting from the excitation of the air flow to the system structures cannot be predicted with this model. In the second step of the study, the model validated with axial fan solutions was applied to the freezer compartment. The aim here is to reveal the air flow distribution in the freezer volume and to identify the regions where turbulence effects increase. In the numerical model, the axial fan was rotated at an operational speed of 1200 rpm and this rotation was achieved by the sliding mesh method. As a result of the analysis, it was seen that the turbulence formation started at the wing tips as observed in the solo fan analyses, and the vortices coming out of the trailing edge tips were especially concentrated in the region between the upper wall of the freezer volume and the upper two drawers. In addition, a turbulent area was detected at the bottom of the evaporator cover (which is the fan suction area). As a result of the hybrid aeroacoustic model solution, the sound pressure data collected from 1 meter away from the front, rear and side surfaces of the freezer and the sound pressure data collected from the same locations in the full anechoic acoustic room were compared. When the total sound pressure in the range of 10-10000 Hz is compared, it is seen that there is a difference of 3-7 dBA between the numerical model and the experimental results. As a result of the investigations of the axial fan in the solo and freezer volume, three different freezer models have been proposed to improve air flow, reduce turbulence and reduce the resulting noise caused by air flow. In the fist suggested model, the bottom part of the evaporator cover has changed and the acostic propagation has decreased 0.24 dBA at 1200 rpm rotational speed. The position of the axial fan and its distance from the structures in the suction and discharge directions are the parameters affecting the acoustic propagation. In the second model, it is aimed to provide acoustic gain by changing the fan position. In this context, the fan was moved on the shaft by 5 mm and brought closer to the blowing region. With this modification, total sound power level was decreased 2.18 dBA. The final model is the superposition of the first two models. Here, it was aimed to see the combined effect of two mentioned model. At 1200 rpm rotational speed, 3.27 dBA gain was achived by the third model.
Aerodynamic and structural optimization a male class unmanned aerial vehicle wing with genetic algorithm

(Graduate School, 2023) Ün, Kağan ; Yıldız, Kaan ; 511191209 ; Aeronautical and Astronautical Engineering Programme

In this thesis, a genetic algorithm based airfoil, planform and wing structure design is utilized for a 1500-2000 kg class fixed wing reconnaissance MALE (Medium Altitude Long Endurance) type UAV (Unmanned Aerial Vehicle). Due to their mission descriptions, these UAV are generally designed to have long range and high loiter time. For these reasons, an aircraft to be designed for observation purposes must have high aerodynamic efficiency and high fuel load ratios (low empty mass) to maximize the range and loiter time. To achieve high aerodynamic efficiency, these aircraft have high wing span ratios, and their wings contain specially designed airfoils. These airfoils are generally designed to have high lift-to-drag ratios. On the other hand, due to long wing structures of high-span aircraft, the wings of such aircraft have high bending forces especially during maneuvers. Therefore, design of airfoils of such aircraft have a compromise between fitting the ideal lightweight inner skeleton for the wing and providing ideal aerodynamics. For these reasons, the wing profile located at the root of the wing is generally chosen to be thicker than the profile used at the tips of the same wing. To simplify the design process, each airfoil are constructed from two Bezier curves that create centreline and thickness distribution of airfoils. Control points of the Bezier curves are the most of the input parameters of the genetic algorithm program. From sets of control points, the airfoils are created. Then, the airfoils are analysed in XFoil by the interface of the function made in MATLAB. After analysis of airfoils, a planform that uses these root-tip airfoils is tested for having sufficiently high lift and low drag for the cruise altitude, cruise speed and cruise power. Then the airfoilplanform combination that pass the basic requirements are sorted by their maximum lift-to-drag ratios. The airfoil-planform combinations with higher maximum lift-todrag ratios are selected for creating the next generation, and the cycle continues. When the maximum number of generations are achieved, the best airfoil-planform combination of the last generation is selected as the best candidate. The fitness criterion of this first phase is the lift-to-drag ratio of the airfoil-planform combination. After the winner airfoil-planform combination is created, inner structure optimization process for the wing begins. Inner structure of the planform consists of four ribs and a twin-box spar structure made of 7068 aluminium alloy, the strongest commercial aluminium alloy available. The lift force and torsion moment of each wing segment is transferred to the spar by the ribs of the wing. The cross section of the spar consists of two closed cells with the support of four stiffeners and eight flanges. Vertical walls have thickness of 2.5 mm, while upper and lower walls have thickness of 1.5 mm. The flanges have cross section value of 400 mm2 , and are set to the upper and lower ends of vertical walls, filling the corners of each cells. The stiffeners have cross section value of 200 mm2 , and are set to the middle of the upper and lower walls, in between the vertical walls. While the stiffeners and flanges carry the tensile and compressive loads, the walls primarily carry the shear loads. To simplify the structural analysis, several assumptions are made. The main spar is assumed as a serial combination of smaller spar sections with constant cross sections. The stresses on cross sections are analysed with structural idealization method. The cross section is assumed as collections of idealized shear force carrying panels and normal force carrying area members called booms. At first, the effective areas at the positions of booms are found by adding effective boom area of wall sections to the actual stiffener or flange areas. From the effective areas and positions of each boom, area moments of inertia and bending moment centre are found. From area moments of inertia of the section and applied bending moment, the compressive and tensile forces of each boom are calculated. After the calculation of tensile and compressive forces, shear forces on the walls are calculated from the area and wall thicknesses of each cell, torsion moment inflicted on section, and compressive and tensile forces of each boom. After the stress calculations are made for each section, a selection process is carried out to control the stresses on the cross sections on each wing rib. If the stresses on the wing at any point is larger than safe limits, the spar is considered as infeasible specimen. Otherwise the fitness value for the spar is compared to the other successful specimens and the fittest specimen is chosen for each generation to create new specimens. The weight of the spar is the fitness value for the second phase. At the end of the process, the ideal cross section is obtained and the program finishes working.
Aktif kontrol uygulamalarında firar kenarı emiş yüzeyi manipülasyonunun akış gürültüsü üzerindeki etkilerinin incelenmesi

(Lisansüstü Eğitim Enstitüsü, 2024-02-05) Toksavul, Atila ; Zafer, Baha ; 511191185 ; Uçak ve Uzay Mühendisliği

Hava araçları ve rüzgâr türbinlerinin aerodinamik ve aeroakustik özelliklerinin iyileştirilmesi, performansları ve işleyişleri açısından önemli bir yer tutar. Havacılık alanında gerçekleşen gelişmeler ile birlikte, uçan yolcu sayısı her geçen gün artmaktadır. Benzer bir şekilde; hızla artan şehirleşme, beraberinde yüksek enerji ihtiyacını getirmektedir. Dolayısıyla, rüzgâr çiftlikleri gibi alternatif enerji kaynaklarının kullanımı her geçen gün artmaktadır. Bu sebeple; gürültü kontrolü ve gürültünün önlenmesi, günümüzde giderek önem kazanmaktadır. Bu çalışma kapsamında; aktif akış kontrol tekniklerinin kullanımının, bir kanadın aerodinamik ve aeroakustik performansı üzerindeki etkileri incelenmektedir. Aktif kontrol metodu olarak, firar kenarına yakın bölgelerde emme akışı uygulamasına gidilmiştir. Firar kenarı gürültüsünün azaltılmasında, bahsedilen aktif akış kontrol metodunun etkinliği araştırılmıştır. Bu doğrultuda, oluşturulan farklı senaryolar için Hesaplamalı Akışkanlar Dinamiği (HAD) analizleri gerçekleştirilmiştir. Sayısal verilerin alınması sürecinde, literatürde yaygın olarak kullanılması ve doğrulama verilerine ulaşılabilmesi nedeniyle, bir NACA (National Advisory Committee for Aeronautics) kanat profili olan ve körlenmiş firar kenarına sahip NACA 0012 profili kullanılmıştır. Kullanılan NACA 0012 profili için veter uzunluğu 0.2m, firar kenarı körlüğü (h/δ*) ise 0.196 olarak alınmıştır. Simülasyonlar, ANSYS Fluent yazılımı üzerinde Reynolds-Averaged Navier–Stokes (RANS) ile Large Eddy Simulation (LES) karma modeli olan Stress-Blended Eddy Simulation (SBES) türbülans modeli kullanılarak gerçekleştirilmiştir. Türbülans alt modeli (subgrid-scale) olarak, Smagorinsky-Lilly modeli kullanılmıştır. Basınç- hız bağıntısının sağlanması için Pressure-Implicit with Splitting of Operators (PISO) algoritması kullanılmıştır. Öncelikle, çalışmada kullanılacak HAD analizleri için oluşturulan sayısal modelin geçerliliği teyit edilmiştir. Bu süreçte, doğrulama analizleri gerçekleştirilmiştir. Gerçekleştirilen doğrulama analizleri sırasında herhangi bir aktif kontrol metodu kullanılmamıştır. Fakat, türbülansa bağlı etkilerin daha iyi gözlemlenebilmesi için, NACA 0012 profili üzerinde hücum kenarına yakın bölgeler üzerinde tetikleme akışı uygulanmıştır. Tetikleme akışı için kullanılan değerler, literatürde yer alan akış verilerini baz almaktadır ve deneysel sonuçlara yakınsamak için sayısal analizlerde kullanılmak üzere oluşturulmuştur. Sayısal modelin doğrulanması, NACA 0012 profili üzerinden elde edilen basınç değerleri ve akustik verilerin literatür ile karşılaştırılması ile sağlanmıştır. Bu süreçte, veter uzunluğuna göre Reynolds sayısı 400000 olarak alınmıştır. Profil üzerinden elde edilen negatif basınç katsayıları, literatür ile karşılaştırılmıştır. Akustik veriler ise, firar kenarına yakın konumlar üzerinden hesaplanarak referans alınan değerler ile karşılaştırılmıştır. Elde edilen sonuçlar neticesinde, kullanılan sınır şartlarının ve oluşturulan sayısal modelin geçerliliği sağlanmıştır. Doğrulama analizleri sonrasında, firar kenarını referans alan doğrultusal gürültü dağılımları, hesaplanmıştır. Elde edilen sonuçlar üzerinde gürültünün kuadrupol etkileri yeterince gözlemlenemediğinden dolayı serbest akış hızı arttırılmıştır. Veter uzunluğuna göre 530000 Reynolds sayısına sahip akış üzerinde aktif gürültü kontrol metotları uygulanmıştır. Farklı aktif kontrol parametrelerinin profil üzerinde uygulandığı senaryolar oluşturulmuş, bulunan sonuçlar birbiri ile karşılaştırılmıştır. Senaryolar, literatürde yer alan aktif gürültü kontrol parametresi baz alınarak oluşturulmuştur. Parametre, aktif kontrol yüzeyinin uzunluğu ve kontrol akışının hızı ile doğru, serbest akış hızı ile ters orantılıdır. Dolayısıyla; serbest akış koşulları değişmediği sürece, birbirine denk kontrol akışı debilerine sahip senaryoların sahip olduğu kontrol parametreleri de birbirine eşit olacaktır. Çalışmada, kontrol yüzeyinin manipüle edilmesi ile oluşturulan farklı senaryolar için gürültü değerleri incelenmiş, birbirine denk kontrol parametrelerine sahip senaryolar üzerinden elde edilen sonuçlar karşılaştırılmıştır. HAD analizleri için oluşturulan modellerde kontrol parametreleri; 0.268, 0.537, 1.611 ve 3.222 olarak belirlenmiştir. Her bir kontrol parametresi için, NACA 0012 profili yüzeyine, birbirine denk merkezli ancak farklı uzunluklarda kontrol yüzeylerine sahip üç farklı durumda aktif kontrol senaryosu tanımlanmıştır. Kontrol parametresinin 0.268 olarak tanımlandığı durumda oluşturulan tüm senaryolar için gerçekleştirilen analizlerde, firar kenarına göre 90° ve 270° doğrultularında gürültü iyileştirme gözlemlenmiştir. Kontrol yüzeyinin daraltıldığı senaryoda bu etkiler artmış, tüm açısal doğrultularda efektif bir gürültü iyileştirme elde edilmiştir. Kontrol parametresinin 0.537 olarak tanımlandığı durumda ise farklı olarak, kontrol yüzeyinin genişletildiği senaryoda etkin bir gürültü iyileştirme gözlemlenmiştir. Tüm doğrultularda efektif olarak gürültü iyileştirmenin gözlendiği senaryolarda; 0° doğrultusunda yani iz bölgesinde gözlenen gürültü miktarındaki azalmanın, diğer doğrultularda gözlenen değerlere göre daha düşük olduğu saptanmıştır. Kontrol parametrelerinin 1.611 ve 3.222 olarak belirlendiği senaryolar ise, elde edilen verilerin doğrulaması, dipol ve kuadrupol etkilerin daha iyi gözlenebilmesi için analiz edilmiştir. Kontrol parametresinin 1.611 olarak belirlendiği senaryolarda iz bölgesi gürültüsünün önlenmesinde etkili sonuçlar elde edilmiştir, ancak bu durum 90° ve 270° doğrultularında gürültü önleme performansında düşüş yaratmıştır. Kontrol parametresinin 3.222 olarak belirlendiği senaryolarda ise ölçüm alınan tüm doğrultular üzerinde gürültü önleme performansında düşüş gözlemlenmiştir.
Analysis of aircraft landing gear brake induced vibrations

(Graduate School, 2023-01-23) Altınbağ, Öner ; Balkan, Demet ; 511191131 ; Aeronautics and Astronautics Engineering

Today, aviation systems are the product of more than 100 years of work. The most groundbreaking process in these studies was experienced during the cold war years. The achievements of many engineering activities today are based on the knowledge gained in these years. Some major problems have been completely resolved in this progress, and some of them still continue to be active problems. The landing gear system is always critical to aircraft and is the engineering solution for almost all functions on the ground. In recent years engineers have been trying to optimize previous achievements within the framework of weight reduction, reliability, integration, energy consumption, noise reduction, cost reduction, and maintenance activities. One of the most important problems related to landing gear systems from the past to the present is the vibration problem, which we can examine under noise reduction. In this study, the causes of vibrations originating from the landing gear braking system were examined together with previous studies in the literature. A comparative approach to brake-induced vibrations, which is still seen as a problem today, has been sought as a solution using today's tools. In this context, the parameters required for an aircraft landing gear model were calculated with the preliminary design activities used in the literature and industry. With these calculations, a model was created using MSC ADAMS software. Tire models in multibody dynamics simulations for vehicle dynamics were examined. As a result, the most suitable tire model was selected for the scope of the study. The parameters of the relevant tire model have been modified from the result of the tire sizing calculations. Two different vibration frequencies were investigated under four different longitudinal velocity conditions in order to make a valuable comparison. The results obtained from the model were compared and interpreted by using the previous studies from the literature.
Analysis of bird strike on metallic panels

(Graduate School, 2023-06-15) Çayhan, Kenan ; Balkan, Demet ; 511201133 ; Aeronautics and Astronautics Engineering

This thesis investigates the phenomenon of bird strikes, using a combination of literature analysis, statistical analysis, and theoretical models. The study focuses on the potential damage that bird strikes can cause to various parts of an aircraft, which are wind-facing components such as wings, stabilizers, engines, and windshields. The variety of possible outcomes from a bird strike poses a significant threat to aviation safety, as bird strikes account for 90% of Foreign Object Damage (FOD) incidents. As a result, aviation regulations require aircraft to meet specific levels of bird strike tolerance for critical components, and there are a number of certification requirements that airplanes must meet to be regarded safe to fly. To investigate the bird strikes on aircraft, the study uses numerical models, including the Smooth Particle Hydrodynamics (SPH) model, which was used to simulate sandwich plate bird impact experiments. The study concludes that the SPH model may be useful for finite element bird strike case analyses, which can help to improve aviation safety by identifying potential vulnerabilities and developing effective prevention measures. When using a new numerical approach, it is important to compare the results to experimental data to ensure that the simulation accurately reflects reality. Many research studies have included both numerical simulations and experimental data to understand how well the simulation corresponds to real-world scenarios. Experimental studies have traditionally guided aircraft designers in creating structures that are tough enough to withstand bird strikes. However, as aircraft components have become more complex, it has become necessary to develop bird strike simulation programs to design aircraft parts that are both airworthy and can be produced quickly and economically. Furthermore, the optimization process typically involves many iterative steps, which makes computer-based analyses more efficient and cheaper than experiments. However, conducting experiments with real birds, which are often dead or drugged chickens, presents a number of issues. The reproducibility of experiments, the health of researchers, and the availability of suitable bird models are all concerns. Real bird torsos vary greatly, making it difficult to obtain consistent results. While certification regulations only define the mass properties of the bird, different bird species have different densities, leading to variations in pressure loads between tests. As a result of these difficulties, researchers have begun using substitute bird materials instead of real birds. Advancements in computer technology have led to the development of cheaper and more advanced finite element software since the 1980s. This has allowed scientists to analyze bird strikes numerically due to the low cost, speed, and repeatability of the analyses. Various substitute bird models have been investigated in studies, and results have been compared with experimental data. The simple cylinder geometry is still a valuable approach to compare simulation results with experimental data. Different geometries such as spheres, cylinders with flat or hemispherical ends, and ellipsoids may also be used in simulations. When birds are struck at high speeds, their behavior is different from that of a simple elastic solid, and it is the responsibility of scientists and engineers to study the behavior of bird materials both theoretically and experimentally. Statistical data related to bird strikes is provided in the thesis, and it is emphasized that front-facing components of aircraft are the most critical as they are most likely to encounter a direct bird strike. The most frequently struck parts of an aircraft are the fuselage, nose, radome, windshield, wing, rotor, and jet engine. Approximately 70% of bird strikes occur at altitudes between zero and 152 meters, which is primarily during takeoff and landing. This information is useful in avoiding bird strike accidents. As the altitude of an aircraft increases, the natural habitats of birds become further from the plane. The velocity of the projectile has a significant impact on how it responds upon impact. The behaviour of the projectile can be divided into five categories based on the internal stresses it experiences: elastic impact, plastic impact, hydrodynamic impact, sonic impact, and explosive impact. Elastic impact occurs when the projectile material strength is well above the internal stresses caused by the low speeds and accelerations, resulting in the projectile bouncing back from the surface. As the impactor velocity increases, the projectile enters the plastic behavior region, yet the velocity is still low enough to maintain fluid-like flow behavior, causing the bird to spread in every direction parallel to the plate, and the load to expand to a larger area. The theory behind bird strike at velocities that cause the bird to act in the hydrodynamic region is investigated. When the impactor with the initial velocity hits a surface, materials in contact with the rigid plate would immediately come to rest, generating a shock wave with velocity normal to the plate and towards the impactor body. There would be a significant pressure gradient at the outer surface because there is shock load pressure on the inner side and free surface pressure on the outer side. Soft objects impacted at high velocities behave differently than at low velocities, such that even elastic solids behave like liquids. However, testing with real birds can yield scattered data and it is not ethical to kill animals for scientific purposes. Gelatine has been found to be a suitable artificial substitute material with uniform characteristics and can be shaped into simple geometries such as cylinders and spheres for easy handling. Finite element programs offer various solution methods for bird strike simulations. Lagrangian method involves nodes attached to the material while Eulerian method uses fixed nodes in a defined space where material flows through it. Arbitrary Lagrangian Eulerian method is another option that allows for the defined space to change with the material flow, leading to faster computation time. Additionally, the meshless method called smooth particle hydrodynamics allows for particles to move freely without mass distortion. Various basic shapes of birds can be examined for bird strike impacts, including a cylinder, a cylinder with hemispherical ends, an ellipsoid, or a sphere. For a bird with a mass of 1.8 kg and specific geometric parameters, the density of the bird can be determined to be 900 kilograms per cubic meter. Conversely, by using a standard density of 950 kilograms per cubic meter and entering the mass of the bird, a specific volume value can be determined and used to specify the bird's geometry. Honeycomb materials provide stiffness to the structure while not adding too much mass. Hence, honeycombs are a kind of deformable shock absorbers that is widely used in the aircraft industry. In the reference tests, they used single and double core honeycomb sandwich metal plates as specimens under bird strike. They made a correlation between test results and simulation results which can be beneficial. Modelling the material of honeycomb in LS-DYNA has a number of challenges. Firstly, honeycomb has a complex geometry which is expensive to model and simulate with shell elements. Therefore, its effective behavior can be modelled under homogenized solid elements. Out of plane stress strain curve up to crushing was given at reference. Which can be inserted as a stress strain curve to the solid elements. Particle node quantity for the bird impactor and element number for the aluminum sheets and honeycomb is limited with the computer power. Therefore, node numbers are generally about 20519 for the bird material. The simulations provide spatial displacement values and nominal strain curve values that are generally similar to the experimental results. However, there are slight differences, which may be due to errors in both the simulations and the tests. Overall, the strain values align well with the experimental data for both simulations. Therefore, the SPH method can be effectively used to simulate bird strikes on honeycomb sandwich plates, which is advantageous since experimental studies can be time-consuming and costly, especially in the initial design phase of aerospace vehicles.
Analytical investigation of quasi-aeroservoelastic behaviour of an aircraft spoiler

(Graduate School, 2022) Kurtiş, Yiğit ; Mecitoğlu, Zahit ; Muğan, Ata ; 777765 ; Aeronautics and Astronautics Engineering Program

The application of science and mathematics to solve problems is called as engineering. In most of engineering process, accuracy is directly dependent on cost which can be defined as function of time and money. In the problem solving processes, there are a lot of assumptions in exchange for accuracy in order to reduce cost and find more solutions in a short time. Reducing solution time provides ability to enhance problem solving capabilities by increasing number of ways to solve problems, finding different sources of problems or optimizing solution methods. At the end, exact solution may not be reached, but more related problems can be solved with approximate solutions in limited time. With advanced technology in aviation industry, accurate designs are more important than before due to desire for better performance. In order to increase accuracy, research and development studies are performed, such as analytical formulizations and tests. Owing to high cost and long durations of test operations, analytical solutions are preferred to be supported if possible. Especially for aircraft design, due to safety consideration and aim for lightweight designs, designers have to balance time, weight and cost without any penalty for safety. In this condition, analytical solutions helps to reduce solution time for lightweight designs and create extra time for optimization studies. In this study, behavior of spoiler structures are investigated for desired deflection angle under external loads by means of analytical solutions. Spoiler is a control surface which can create drag and lift for aircraft. Spoiler structures have been implemented to aircrafts in order to improve control, especially while rolling, landing and braking. One of the main objectives of a spoiler structure is to increase drag for landing and braking applications. Additionally, spoilers can be used to increase roll rate for acrobatic or trainer aircrafts. Under aerodynamic load, as all structures deform, spoiler structures show a deformation. It affects spoiler deflection mechanism because points of mechanism changes when spoiler deformation occurs. In this case, spoiler rotates back towards to its original position where back rotation angle is usually not able to be considered in mechanism design. This condition creates dwindle for effectivity of spoiler surface which means reducing performance of aircraft. In this thesis, an analytical formulation study is performed in order to foresee back rotation angles of spoiler structures and gain ability to design mechanism for more convenient deflection angles for spoilers under aerodynamic loads. Result curves are created by curve fitting method in order to monitor and compare behavior of both analytical method analyses and finite element method analyses. Error functions are defined and calculated to find out tendency difference between analytical method and finite element method analyses under changing variables. For realistic deflection angles, the aim of this study is to accomplish accurate analytical results with error percentage below ±15% for back rotation angles and ±2% for final deflection angle compared to finite element method analyses. In the introduction section, engineering approaches for development studies are explained. Importance of accuracy for engineering application is tried to be stated by support of relations between accuracy and other engineering concerns. These concerns can be expressed with time, cost and other concerns, such as health issues, ethic concerns and safety. Also, scope and purpose of thesis are determined in this section. In the literature review section, spoiler structures and their duties on aircraft are stated. Dimensions are shown with examples and figures from aircraft industry. Grid stiffened spoiler concepts are explained in addition to commonly used structural architectures, such as composite and metal builtup structures.
Attitude estimation and reaction wheels based control of an earth-pointing small satellite

(Graduate School, 2024-06-24) Gürsoy, Hakan ; Hacızade, Cengiz ; 511211123 ; Aeronautics and Astronautics Engineering

A satellite is an artificial object that is sent into orbit around a celestial body, usually the Earth. Depending on their mission design, they can perform various tasks from communication to military. Satellites consist of various subsystems to perform their assigned mission. Attitude Determination and Control Systems (ADCS) is responsible for orienting the satellite in the desired direction and maintaining its orientation in space. It uses various sensors and actuators to measure and control the satellite's attitude. The sensors, such as sun sensors, magnetometers and gyroscopes, provide data about the satellite's current orientation. The measurement data coming from these sensors are processed by various attitude estimation methods. The actuators, like reaction wheels, magnetorquers and thrusters, apply the necessary torques to adjust the satellite's attitude. The ADCS ensures that the satellite's payload is correctly positioned for its mission. If the satellite cannot be brought to the required orientation, it may not be able to fulfil its mission. In this thesis, some of the prominent vector measurement-based attitude estimation methods are compared. To make this comparison, the dynamic and kinematic model of an Earth-pointing satellite is derived and subsequently, this highly nonlinear model is linearized by using the Taylor series method. Following this, mathematical models of the sun sensor and magnetometer were also presented. Real sensor measurements are simulated by adding white noise to these mathematical models. The compared attitude estimation methods can be named as TRIAD, Q-Method, and SVD. The comparison is made through conducting simulations in the Matlab/Simulink environment. Root mean square errors of the estimation methods are computed by comparing their outputs with the output of the deterministic satellite system model. After comparing the attitude estimation methods, a comparison was made between the two types of the optimal controllers. The LQR controller is a feedback control method that aims to minimize a quadratic cost function, which is typically defined in terms of deviations from desired states and control inputs. It provides an optimal solution for controlling the satellite's attitude while considering both system dynamics and control effort. The LQG controller improves the capabilities of LQR by adding a Kalman filter, improving the estimation of system states from noisy sensor measurements. This integration improves the controller's ability to handle uncertainties and disturbances, making it more suitable for real-world satellite applications where sensor data may be prone to noise or inaccuracies. Reaction wheels are selected as the actuators. Reaction wheels are based on the principle of conservation of angular momentum. When it is necessary to change the orientation of the satellite, reaction wheels start to rotate according to the control signal; causing a change in angular momentum. The satellite body produces an angular velocity in the opposite direction to preserve the total momentum. In this way, the orientation of the satellite can be controlled. Although three reaction wheels are sufficient for attitude control in three axes, satellites generally use reaction wheel configurations consisting of four wheels. An extra fourth wheel creates actuator redundancy. If one of the reaction wheels fails, the other remaining three reaction wheels can still complete the mission. The internal dynamics of the reaction wheels are modelled as a DC motor on Simulink. There are various disturbance torques that affect the satellite along its orbit. For example, gravity-gradient torque comes from the Earth's unequal gravitational force acting on different parts of the satellite. The parts of the satellite that are closer to the Earth are exposed to more gravitational force. This imbalance creates a torque on the satellite. Disturbance torques like gravity-gradient torque however can also be used to keep the satellite stable. In this thesis, the gravity-gradient stability behaviour of the simulated satellite was also examined.
Autonomous heading control of a fixed-wing aircraft with deep reinforcement learning

(Graduate School, 2022) Sarıgül, Fatih Ahmet ; Beyazit, İsmail ; 771377 ; Aeronautics and Astronautics Engineering Programme

Autonomous control has become more reachable with recent advancements and new techniques such as deep reinforcement learning (DRL) and is getting more and more popular in every field including flight control. Autonomous flight is an important trait to attain for an aircraft because it provides to get rid of external involvement in control and it has the potential to excel in human skills. In fully autonomous flight, the aircraft is needed to complete its all tasks without any human involvement. However, instead of fully autonomous flight, this work focuses only on heading control because making the heading control autonomous by a learning algorithm means that it is possible to make other flight tasks autonomous by using a similar learning algorithm. Therefore, in this work, devising a learning algorithm for autonomous heading control is the main goal of the work. In this work, it is desired to attain autonomous control for a case that demands a complex environment and dynamics like the heading maneuvers of a fixed-wing aircraft. However, because the focus of the work is mainly at the algorithmic level, it is not a good idea to dive into the complex environment and dynamics directly. Therefore, firstly the Dubins model is used in this work to represent the fixed-wing aircraft while testing the learning algorithm. Dubins model can be simply considered as just a point that has a heading and a velocity. It can be implemented in both 2D and 3D environments easily and it can represent a car driving or an aircraft flight in a very basic manner. However, with some constraints and some addition to its dynamics, it can be converted easily into a good representation of a fixed-wing aircraft. This new representation is much simpler than a fully described dynamics of a fixed-wing aircraft but it is still a good representation to see how the learning algorithm works. Therefore, in this work, a simplified fixed-wing aircraft model is obtained from a 3D Dubins Airplane model and the learning algorithm is tested in this simplified model. Implementation of the learning algorithm to a fully described 6-degree of freedom dynamics is left for future works. It seems that the state-of-art DRL algorithms can provide a solution for the autonomous heading control of a fixed-wing aircraft. So, finding the most appropriate state-of-art algorithm and implementing it to the problem can offer a solution. But, in this work, it is not the way that is followed. Instead of jumping directly to the state-of-art methods, this work starts with a basic learning algorithm, and it is tested in a simple environment, after getting satisfactory results at this level, the environment is rendered more complex status and the algorithm is made more advanced to deal with new conditions. In this way, it can be seen why the previous methods lack and what is needed to make learning possible in this new condition. By repeating this process, it is aimed at obtaining a DRL algorithm to solve the problem while having the opportunity to make improvements in the algorithm at every level. In this work, the DRL algorithm is obtained by using the aforementioned technique for heading control. This DRL algorithm consists of a combination of DQN and Actor-Critic methods. It meets the requirement of dealing with continuous state and action spaces and it has some unique approaches which do not exist in other algorithms. The new algorithm that has been obtained in this work is tested on the Dubins model and simplified model to see its validity and whether it can be used for more complex tasks and dynamics. The promising results show that the algorithm can be enhanced to deal with other flight tasks also, and it may offer solutions to complex real-world problems.
Beşinci nesil muharip uçaklarda kullanılan kompozit malzemelerin bazı mekanik özelliklerinin ve kesilme parametrelerinin incelenmesi

(Lisansüstü Eğitim Enstitüsü, 2024-07-18) Ayan Sayın, Elif Gizem ; Caferov, Elbrus M. ; 511221109 ; Uçak Mühendisliği

Günümüz havacılık endüstrisi, hızlı ve sürekli gelişen dinamik yapısıyla dikkat çekmektedir. Bu sektörde önemli bir rol oynayan muharip uçaklar, hız, manevra kabiliyeti ve teknolojik üstünlük açısından sınırları zorlamaktadır. Muharip uçakların gelişiminde özellikle beşinci nesil uçaklar, öncü teknolojilerle donatılarak dikkat çekmektedir. Beşinci nesil muharip uçaklar, gelişmiş aviyonik sistemler, ileri radar teknolojileri ve kompozit malzemelerin kullanımıyla tanımlanmaktadır. Kompozit malzemeler, hafiflik, yüksek mukavemet ve dayanıklılık gibi üstün özellikleri sayesinde havacılık endüstrisinde önemli bir yer edinmiştir. Bu malzemeler, uçak performansını artırmak ve yakıt verimliliğini optimize etmek amacıyla kullanılırken, geleneksel metallerin yerini giderek daha fazla almaktadır. Ancak, kompozit malzemelerin avantajları kadar, bu malzemelerin karmaşıklığını ve davranışlarını anlamak da büyük önem taşımaktadır. Bu çalışma, beşinci nesil muharip uçaklarda yaygın olarak kullanılan kompozit malzemelerin uzunluk, kalınlık ve düzen özelliklerinin eğilme gerilmesi üzerindeki etkilerini detaylı bir şekilde incelemektedir. Çalışmanın temel amacı, bu malzemelerin farklı parametreler altında eğilme gerilmesine nasıl tepki verdiğini anlamak ve tasarım sürecinde performanslarını optimize etmek için önemli bir katkı sağlamaktır. Ayrıca, çalışma kapsamında kesme parametrelerinin kompozit malzemeler üzerindeki etkisi daha yakından incelenmektedir. Çalışmada kullanılan bal peteği yapılı kompozit malzemeler yedi farklı kombinasyonda üretilmiş, lamina kompozit malzemeler ise üç farklı kombinasyonda üretilmiştir. Bu malzemeler üzerinde dört noktalı eğilme ve üç noktalı eğilme testleri gerçekleştirilerek performansları karşılaştırılmıştır. Elde edilen sonuçlar, tablolar ve grafikler aracılığıyla detaylı bir şekilde sunulmuştur. Karşılaştırma sonucunda, maksimum kalınlıkta, iki kat karbon fiber ve bir kat Kevlar'dan (aramid) oluşan kompozit malzemenin en yüksek eğilme dayanımını sergilediği gözlemlenmiştir. Ayrıca, sandviç panellerde bal peteği ve karbon fiber kalınlığının artırılmasının eğilme dayanımını olumlu yönde etkilediği tespit edilmiştir. Bu bulgular, tasarım mühendislerine beşinci nesil muharip uçakların malzeme seçimi ve konfigürasyonunda rehberlik edecek önemli bilgiler sunmaktadır. Kompozit malzemelerin davranış ve özelliklerini kapsamlı bir şekilde anlayarak, mühendisler bilinçli kararlar alabilir ve uçak performansını ve dayanıklılığını artırarak havacılık endüstrisinin gelişimine katkıda bulunabilirler. Özellikle bal peteği ve lamina kompozit malzemeler üzerinde yapılan bu detaylı incelemeler, mühendislik tasarım süreçlerine ışık tutmakta ve yeni nesil muharip uçakların geliştirilmesinde kritik rol oynamaktadır.Çalışmada kullanılan dört noktalı eğilme ve üç noktalı eğilme testleri, malzemelerin eğilme dayanımı ve performansı hakkında önemli veriler sağlamıştır. Elde edilen sonuçlar, kompozit malzemelerin yapısal tasarımda nasıl optimize edilebileceğine dair önemli ipuçları sunmaktadır. Özellikle karbon fiber ve Kevlar kombinasyonlarının üstün performansı, bu malzemelerin yeni nesil muharip uçak tasarımlarında daha yaygın bir şekilde kullanılabileceğini göstermektedir. Sonuç olarak, bu çalışma, beşinci nesil muharip uçakların malzeme seçimi ve tasarımında önemli bir rehber olma niteliği taşımaktadır. Kompozit malzemelerin farklı parametreler altında nasıl davrandığını anlamak, mühendislerin daha sağlam, hafif ve verimli uçaklar tasarlamalarına olanak tanıyacaktır. Buda, havacılık endüstrisinin gelişimine ve teknolojik ilerlemelere önemli katkılar sağlayacaktır. Çalışmanın bulguları, sadece akademik alanda değil, aynı zamanda pratik mühendislik uygulamalarında da geniş bir kullanım alanı bulacaktır. Bu sayede, geleceğin muharip uçaklarının daha güvenilir ve etkili bir şekilde tasarlanması mümkün olacaktır.
Bir haberleşme uydusunda faydalı yük panel tasarımı ve kalifikasyonu

(Lisansüstü Eğitim Enstitüsü, 2023-06-20) Canbolat, Uğur ; Mecitoğlu, Zahit ; 511181140 ; Uçak ve Uzay Mühendisliği

Bu çalışmada, bir haberleşme uydusunun önemli bir yapısal bileşeni olan faydalı yük sandviç panelinin yapısal tasarımı ve kalifikasyon süreci incelenmektedir. İlk bölümde uydu sınıflandırılması, uydu yapısı ve yapısal alt sistemleri hakkında genel bir bilgi verilmiş ve yer durağan Dünya yörüngesinde bulunan bir haberleşme uydusunun özellikleri ve gereksinimleri tanıtılmış ve uydunun fırlatma aşamasında karşılaşacağı statik, dinamik, akustik vb. yüklemelerle alakalı bilgiler verilmiştir. Belirlenen gereksinimlere göre tasarlanmış bir haberleşme uydusunun sonlu elemanlar modeli Hyperworks/Hypermesh ticari yazılımı kullanılarak oluşturulmuştur. Modelleme aşamasında kullanılan metodlar detaylı olarak anlatılmış ve oluşturulmuş sonlu elemanlar matematiksel modelinin kontrolü ECSS standartlarına göre kontrol edilmiştir. Uydu fırlatma yükleri altında faydalı yük paneline etki eden yüklerin türetilme metotları tanıtılmıştır. Bu çalışmada uydu birincil ve ikincil yapılar için kritik yük girdilerinin sinüs titreşiminden kaynaklandığı Ariane 5/6, Proton, Falcon 9 veya New Glenn aday fırlatıcılarının sinüs yük zarfları dikkate alınmıştır. Bu fırlatıcıların boylamsal ve yanal sinüs fırlatma yükleri dikkate alınarak MSC Nastran ticari yazılımı kullanılarak birincil çentik indirgeme analizi yapılmıştır. Birincil çentik indirgeme analizi aşırı yüklemeye (overtesting) maruz kalmaması için yapılan bir çalışmadır. İndirgenmiş sinüs yükleme girdileri için frekans cevap analizi tekrarlanmış, kuzey faydalı yük paneli ve güney faydalı yük paneli üzerinden düzlem dışı ivme dağılımları hesaplanmıştır. Faydalı yük paneli üzerindeki ekipman ağırlık dağılımları ve sinüs ivme cevapları dikkate alınarak düzlem dışı statik eşdeğer yük değerleri elde edilmiştir. Hesaplanan statik eşdeğer yükler için panel seviyesi sonlu elemanlar analizi yapılmış ve güvenlik marjının düşük olduğu bölgeler için dayanımının arttırılması ve sandviç panel ağırlığını hafifletmek amacıyla tasarım kısıtlamaları göz önüne alınarak bal peteği çekirdek yapıda tasarım iyileştirme iterasyonları gerçekleştirilmiştir. Nihai tasarım modelinin üretimi gerçekleştirilmiş ve hesaplanan statik eşdeğer yükler altında sandviç panelin kalifikasyonu için statik test kampanyası planlanmıştır. Gerçek uydu bağlantı arayüzü dikkate alınarak test fikstürü tasarlanmıştır. Test numunesi 137 adet ISO4762-M6-12.9 bağlayıcı ile test fikstürüne sabitlenmiştir. Test öncesi panel seviyesi sonlu elemanlar analiz sonuçları kullanılarak gerinim ölçer ve deplasman ölçer lokasyonları belirlenmiştir. 13 adet rozet tipi gerinim ölçer ve 4 adet yer değiştirme (deplasman) ölçer sensörü yerleştirilmiştir. Test yüklemeleri hidrolik pistonlarla kademeli olarak %25, %50, %75, %90 adımlarıyla kalifikasyon seviyesine ulaşmış ve sonra aynı adımlarla yük boşaltma yapılmıştır. Her adımda test değerleri ile sonlu elemanlar analiz değerleri karşılaştırılmış olası bir hasar durumunda müdahale etmek için izlenmiştir. Test başarım kriterleri kapsamında test numunesi incelenmiştir. Gerinim ölçer ve yer değiştirme (deplasman) ölçer sensörlerinden elde edilen veriler sonlu elemanlar analiz sonuçları ile karşılaştırılmış ve sonuçlar değerlendirilmiştir.
Bir savaş uçağının burun iniş takımı yapısal analizi

(Lisansüstü Eğitim Enstitüsü, 2023-01-23) Aydın, Gözde ; Özkol, İbrahim ; 511191200 ; Uçak ve Uzay Mühendisliği

Hiç şüphesiz uçak tasarımında uçağın her bir komponent ayrı bir mühendislik süreci ve ciddi bir zaman gerektirmektedir. Uçaklarda iniş takımı önemli bir ana mekanik sistemdir. Bu tez çalışmasında da bir savaş uçağının burun iniş takımı tasarımı gerçekleştirilerek yapısal analizi yapılmıştır. Tez çalışması süresince, birçok kaynak incelenmiştir ve uzun bir literatür araştırma süreci gerçekleştirilmiştir. Havacılık endüstrisinde yüksek dayanımlı ve hafif bir yapı tasarlamak en kritik parametrelerdendir. İniş takımları uçakların toplam ağırlığının yaklaşık %6' sını oluşturur. Yani uçak ağırlığının büyük bir kısmını oluşturur. Dayanım/ağırlık oranı yüksek iniş takımı tasarlamak en önemli tasarım gerekliliğidir. İniş takımları, iniş ve kalkış sırasında uçağa gelen dinamik ve statik yüklere maruz kalır. Bu nedenle iniş takımı sisteminin bu yüklemelere karşı dayanımlı bir yapıya sahip olması gerekir. Yüklere dayanamadığı takdirde iniş takımı ve uçakta ciddi yapısal hasarlar meydana gelebilir. Havacılık tarihinden bu yana birçok farklı çeşitte iniş takımları tasarlanmıştır. İlk başta tasarımlarda sabit iniş takımları kullanılırken zaman içerisinde bu tip iniş takımlarının aerodinamik açıdan dezavantajlı olduğu görülmüştür. Uçaklarda daha yüksek hız ve daha uzun havada kalma süresi gibi isterleri karşılayabilmek için katlanabilir iniş takımları tasarlanmıştır. Daha kompleks bir yapı olmasına karşın uçaklarda performans isterleri de göz önüne alındığında katlanabilir iniş takımlarının kullanımı zamanla yaygınlaşmıştır. İniş takımı tipine karar verildikten sonra ana ve burun iniş takımının konumuna karar verilirken, ağırlık merkezinin konumu göz önüne alınarak uçağın yerde hareketi, devrilmemesi, yan rüzgar etkisini azaltması, iniş ve kalkış sırasında manevra kabiliyetine izin vermesi sağlanmalıdır. Öncelikle, iniş takımı analizi için bir tasarım hazırlanmıştır. Bu tasarım için bilinmesi gereken belli parametreler vardır. Bu parametreler uçağın ağırlık merkezi, yerden yüksekliği, ortalama veter uzunluğu, iniş takımları arasındaki mesafe gibi sıralanabilir. Literatürdeki savaş uçaklarının bir çoğu incelenerek bu parametreler ile ilgili veriler toplanmıştır. Ortalama bir değer seçilerek kavramsal tasarım için gerekli bilgiler elde edilmiştir. Böylelikle iniş takımının uçağın ağırlık merkezine göre yerleşimi yapılmıştır. Daha sonra, uçak yerleşimine göre iniş takımlarına gelen yüklemeler hesaplanmıştır. Yük hesaplamalarında literatürdeki kitaplardan faydalanılmıştır. Yüklere göre lastik boyutuu, amortisör stroğu ve dikme çapı belirlenmiştir. Parça çizimleri ve montajda Siemens NX programı kullanılmıştır. Ayrıntılı boyutlandırma ve çizim yapıldıktan sonra kritik yük koşullarını belirleyebilmek için farklı iniş koşullarında iniş takımına gelen üç eksendeki kuvvetler hesaplanmıştır. Farklı kuvvet ve doğrultularda en kiritik üç koşul seçilerek analizler bu iniş koşullarında gerçekleştirilmiştir. Sonlu elemalar yöntemi ile burun iniş takımı ANSYS Workbench programı kullanılarak analiz edilmiştir. Yapısal analizi gerçekleştirilen iniş takımında malzeme değişikliği yapılarak yapıların Von-Mises gerilme ve deformasyon değerleri elde edilmiştir. Yapıların maruz kaldığı yüklemeler yüksek olduğu için komponentlerde yüksek gerilmeler görülmüştür. Bu nedenle, iniş takımlarında malzeme seçilirken yüksek dayanım ve uzun ömre sahip olması önemlidir. Son olarak, yapılan analiz sonuçlarına göre parçaların ağırlıkları, deformasyon miktarları ve dayanımları karşılaştırılmıştır. Bu sonuçlar doğrultusunda tasarım kriterleri de göz önüne alınarak malzeme seçimi yapabilir veya tasarımda değişiklik kararı alınabilir. Bu şekilde yapılan analizler serisi ile optimum bir burun iniş takımı tasarımına ulaşmak mümkündür.
Blast yükü altındaki sandviç kompozit panellerin mekanik davranışının sayısal ve deneysel olarak hesaplanması

(Lisansüstü Eğitim Enstitüsü, 2023) Sertaş, Yunus Emre ; Balkan, Demet ; 807215 ; Uçak ve Uzay Mühendisliği Bilim Dalı

Kompozit yapıyı, iki veya daha fazla farklı malzemenin istenen özelliklerini ön plana çıkararak bir araya getirilmesi olarak tarif edersek, çok eski çağlarda dahi kompozit malzemelerin kullanıldığını söyleyebiliriz. Kompozit malzemeler günümüzdeki popülaritesini ve yaygın kullanımını yeni keşfedilmiş olmasına değil malzeme ve üretim teknolojilerinin gelişmiş olmasına borçludur. Havacılık ve uzay sektörünün bu gelişmelerde başı çektiği göz önünde bulundurulduğunda; kompozit malzemelerin ileri seviye ihtiyaçlara çözüm ürettiği rahatlıkla söylenebilir. Çünkü daha hafif ve daha sağlam yapılar hava ve uzay araçlarının birincil ihtiyacıdır. Bu ihtiyacı karşılamak için yapısal tasarım mühendisleri yük akışının zayıf olduğu yerlerden malzeme eksiltirken, yükün fazla olduğu yerlere malzeme takviyesi yapmaktadırlar. Geleneksel malzemeler ve üretim metodları bu yöntemi bir yere kadar desteklemektedir. Kompozit malzeme teknolojisinin malzemenin iç yapısına müdahale etme ve daha küçük ölçekte farklı malzeme kullanım imkanı mühendislerin optimum tasarıma erişmesine yardımcı olmaktadır. Her ne kadar günümüzde en yaygın olarak havacılık ve uzay sektöründe kullanılsa da kompozit yapılar, gemicilik, inşaat, otomotiv vb. sektörlerde de kullanılmaktadır. Kompozit yapıların bir alt elemanı olarak ifade edebileceğimiz sandviç kompozit yapılar, geleneksel yapılara göre daha yüksek dayanıma sahip ve hafif olması, bunun yanı sıra monolitik kompozit yapılara göre daha yüksek burkulma direncine sahip olmasından ötürü havacılık sektöründe gün geçtikçe daha yaygın kullanım alanı bulmuştur. Temelde iki yüzey levhası arasına görece oldukça hafif çekirdek malzemenin yerleştirilmesiyle elde edilen kompozit yapılar; yükleme durumunda bir I kirişi ile çok benzer davranış gösterir. Yüzey levhaları çekme ve basma yükünü karşılarken, çekirdek kayma yüküne karşı direnç gösterir. Düzleme dik yüklemelerde ise I kirişe göre daha homojen bir dayanıma sahiptir. Sandviç yapılarda yüzey levhası olarak en fazla aluminyum alaşımlar ve fiber takviyeli kompozitler tercih edilmektedir. Çekirdek malzemesi olarak da petek yapı, köpük veya balsa tercih listelerinin en üst sıralarında yer alır. Havacılık sektöründe karbon fiber takviyeli epoksi levhalar ile balpeteği yapılar yaygın kullanıma sahiptir. Özellikle yüksek burkulma ve yorulma dayanımından ötürü hava araçlarının dış panellerinde kullanılmaktadır ve ses hızının üzerinde hareket eden araçlarda döngüsel olarak yüksek basınç yüklemesine maruz kalmaktadır. Ayrıca yüksek sönümleme kabiliyetinden ötürü iniş takımı tekerlekleri gibi patlama durumunda çevresine blast yüklemesi yapabilecek ekipmanların koruma panellerinde kullanılır. Bu çalışmada blast yükü altındaki karbon fiber takviyeli epoksi levhalara ve balpeteği çekirdek yapısına sahip sandviç kompozit panellerin blast yüklemesi altındaki mekanik davranışının sayısal ve deneysel olarak incelenmesi hedeflenmiştir. Bu xxiv hedefe yönelik olarak karbonfiber kompozit yapılar mekanik davranışı üzerine kapsamlı bir literatür araştırması yapılmış; teorik ve deneysel çalışmalar detaylı olarak incelenmiştir. Ardından katmanlı kompozit yapıların mekaniği üzerine teorik çalışmalar yapılmış ve çalışmada kullanılan temel denklemler elde edilmiştir. Katmanlı kompozit yapılar için geliştirilmiş 2 boyutlu ve 3 boyutlu teoriler ile ilgili bilgi verilmiştir. Sandviç yapılarda kullanılan çekirdek yapılar araştırılmış ve projede incelenecek olan petek çekirdek yapısının mekanik özellikleri detaylı olarak incelenmiştir. Blast yük profilinin elde edilmesi ve yükün ölçeklendirilmesi için literatürde kullanılan yaklaşımlar değerlendirilmiş ve teorik çalışmalarda kullanılacak fonksiyonlara karar verilmiştir. Yapılan teorik çalışmalar sonucunda elde edilen sonuçların deneysel olarak sınanması için hazli hazırda bulunan deney düzeneği basit mesnetli plaka sınır koşullarına göre revize edilmiştir. Literatür araştırması ve sayısal çalışmalar sonucunda çekirdek kalınlığı, çekirdek yoğunluk ve çekirdek tipi sayısal ve deneysel olarak incelenecek parametreler olarak belirlenmiştir. Bu parametrelerin deneysel olarak inceleneceği numunelerin üretileceği malzemeler ve üretim metodlarını belirlemek adına çalışmalar yapılmıştır. Sandviç panellerin yüzey levhalarının kalınlığını belirlemek amacıyla, farklı kalınlıklarda monolitik karbonfiber kompozit plaklar üretilmiş ve F-16 uçağının ön tekerinin lastik basıncı değerindeki blast yükü altında plakların mekanik davranışları plak merkezine yerleştirilen gerinim pulları yardımıyla ölçülmüştür. Deneylerden alınan veriler sonucunda tüm numunelerde 1.472 mm kalınlığında karbonfiber kompozit yüzey levhası kullanımının uygun olduğuna karar verilmiştir. Yüzey levhası kalınlıklarının belirlenmesiyle birlikte asıl deneylerde test edilecek sandviç kompozit plakların üretimine başlanmıştır. Numune geometrileri oluşturulmuş ve istenen malzeme özelliklerinin elde edilmesi için gereken sınır koşullarda üretimler gerçekleştirilmiştir. Üretilen numunelerin her birinin merkezine ve köşesine birer 3-eksenli gerinim pulu yerleştirilmiştir. Numuneler teker teker deney düzeneğine basit destekli sınır koşuluna uygun olacak şekilde yerleştirilmiş; her bir plak üzerine aynı basınç profiline ve şiddete sahip olacak şekilde blast yüklemesi yapılmıştır. Gerinim pulları üzerinden eş zamanlı ve yüksek çözünürlükte alınan datalar bilgisayar ortamında anlamlı verilere dönüştürülmüştür. Veriler incelenirken gerinim pullarının kopması sonucu veya rüzgarın etkisi ile ortaya çıkan bozulmalar değerlendirmelere dahil edilmemiştir. Bu deneylerden elde edilen değerler teorik formüllerden elde edilen değerler ile kıyaslanmış ve deney sonuçlarına dair kaba bir hata hesabı yapılmıştır. İlk tur deneylerinin ardından analiz çalışmaları için gereken plak üzerine etkiyen basınç profilini elde etmek amacıyla üç numune üzerine gerinim pullarına ek olarak basınç sensörü yerleştirilmiştir. Bu üç numune üzerinde önceki deneyler ile aynı sınır koşullarına sahip olacak şekilde blast yüklemesi deneyleri tekrar edilmiştir. Deneylerden alınan veriler vasıtası ile basınç yük profilinin hem zamana hem de konuma bağlı olarak değişimini ifade eden eğri denklemi çıkarılmıştır. ABAQUS programı vasıtası ile deneylerde incelenen plakların üzerine sırasıyla bu eğri denkleminden elde edilen basınç profili uygulanmıştır. Programda plaklar, deneylerde olduğu gibi basit destekli plak sınır koşullarına göre mesnetlenmiştir. Plaklarda kullanılan karbonfiber kompozit ve balpeteği çekirdek yapılarının mekanik özellikleri temin edilen firmaların kalifikasyon deneyleri ile elde ettikleri veri listelerinden çekilmiş ve programda her katman için ayrı ayrı tanımlanmıştır. Deneylerde plaka merkezinden elde edilen gerinim değerleri ile analizlerde aynı noktada elde edilen gerinim değerlerinin zamana bağlı değişim profilleri grafikler yardımı ile kıyaslanmıştır. Aynı zamanda yine ABAQUS programı vasıtası ile üç farklı çekirdek kalınlığına sahip sandviç kompozit plakların ayrı ayrı modal analizleri yapılmış ve doğal frekans değerleri elde edilmiştir. Sayısal ve deneysel çalışmalardan elde edilen veriler kapsamlı olarak değerlendirilmiş ve yorumlanmıştır. Teoride yapılan kabullerin sandviç kompozit plakların mekanik davranışına olan etkisi deneysel çalışmalardan alınan sonuçların ışığında incelenmiş ve hata payları ortaya koyulmuştur. Son olarak da bu çalışma ile blast yükünün sandviç kompozit plaklara olan etkisini incelemeye ve detaylandırmaya yönelik yapılacak olan çalışmalar için tavsiyelerde bulunulmuştur.
Case study on selecting optimal main central cone of low earth orbit satellite

(Graduate School, 2023) Türkoğlu, Semih ; Özdemir, Özge ; 856751 ; Aeronautics and Astronautics Engineering Programme

In this thesis, determination and selection of the most optimum central cone design concept for low orbit satellites of 1000 kg and above from large class satellites is explained. The designs of the cone and cone assembly parts in the study were made with Siemens NX, and the analyzes (modal analysis, static structural analysis, buckling analysis) applied for comparisons were made with Simcenter 3D. The study consists of two stages, in the first stage, a total of 6 cones, 3 different design concepts (semi-monocoque, grid, sandwich) from 2 different materials, Aluminum and CFRP, were compared and two cones were selected after the return-lump study. In the second stage, after the iterations on these two cones, the most optimum central cone design was decided by repeating the return-lump study. In comparisons, mass, stiffness, static strength and buckling strength were determined as criteria.
Cavitation performance improvement of engine cooling pump

(Graduate School, 2025-01-22) Karbal, Doğukan ; Edis, Fırat Oğuz ; 511221106 ; Aeronautical and Astronautical Engineering

As part of this study, the cavitation issue observed at the specific operating point of the water pump used in the cooling system of an off-road vehicle was investigated, and an optimization process was conducted to address the problem. The optimization process consisted of three stages. In the first stage, based on the insights gained from a literature review, different designs for the impeller's blade leading-edge profile were tested. The impact of three different blade leading edge configurations (v0, v1, and v2) on the impeller's cavitation performance was investigated using 3D CFD analysis. The initial analyses were performed in steady-state conditions using ANSYS CFX. Regions on the impeller with pressures below the vapor pressure of the fluid were compared, and the optimal design (v2) was selected. Following the selection of the optimal design, cavitation analyses were conducted on the pump model with the baseline impeller and the pump model with the optimized impeller using ANSYS CFX. These analyses employed the Rayleigh-Plesset cavitation model and were solved under time-dependent conditions. The results demonstrated that the optimized design exhibited superior cavitation performance compared to the baseline design. The optimized impeller was then manufactured and subjected to experimental cavitation tests. These tests revealed that the optimized impeller allowed the pump to operate at 30% lower inlet pressure before cavitation occurred compared to the baseline pump, confirming the significant effectiveness of the optimization process. Subsequently, the final impeller design underwent experimental performance testing, and a validation study was carried out by comparing the experimental results with 3D CFD results. The validation analyses were performed using Siemens StarCCM+ in time-dependent conditions. The comparison between 3D CFD results and experimental data showed a maximum deviation of 3%, confirming the validity of the numerical approach. Performance test results were compared both numerically and experimentally using H-Q curves, which are presented for clarity.
Comprehensive investigation of rotor stall onset for future helicopters

(Graduate School, 2024-06-24) Uçar, Enes ; Özdemir, Özge ; 511201162 ; Aeronautical and Astronautical Engineering

Helicopter main rotor operates in different aerodynamic environment in hover and forward flight conditions. In hover, downwash is almost steady and there is not significant amount of variation in aerodynamic loads. In forward flight, periodic aerodynamic fluctuations occur in local Mach number (the ratio of section speed to the speed of sound), flapping, lagging and feathering dynamics. The complex aerodynamic environment of a rotor in high-speed flight is tightly coupled with the aeroelastic and rotor dynamics characteristics. Since the aerodynamic loads are affected by local Mach number changes, cyclic pitch and cyclic flapping effects, it becomes harder to predict the stall boundaries. As the local blade sections observe high pitch fluctuations during forward flight, separations and loss of lift may occur. Unlike airfoil stall, where thrust typically decreases, in rotor stall, thrust may continue to rise post-stall, but a substantial increase in power demand is seen. Oscillatory loads increase on the rotor hub and pitch links. These force oscillations create undesired vibrations on the airframe affecting the passenger and pilot comfort as well as control effectiveness, handling qualities, fatigue life of critical components, limiting the lift and propulsion capability of the rotor. This boundary is called "Stall Onset". Determining this boundary at the preliminary design phase of a helicopter is crucial since it serves as a key indicator of rotor performance. This thesis investigates helicopter main rotor stall onset and the factors influencing the onset boundary such as fuselage parasite drag and lift. Comprehensive modeling is employed to accurately determine rotor stall onset. CAMRAD II (Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics) tool is utilized to structurally and aerodynamically model UH-60A Black Hawk main rotor. CAMRAD II, widely recognized in the aerospace industry, integrates multibody dynamics, nonlinear finite elements, structural dynamics and aerodynamics to perform nonlinear dynamic and static analyses. The UH-60A isolated main rotor model generated in CAMRAD II has 21 aerodynamic panels for each blade, 20 stations to define structural parameters and 40 stations to define blade mass properties. The geometry of rotor pitch control system (pitch links), rotor control system and non-linear lag damper are also integrated into the rotor model. The model is verified through comparisons with wind tunnel test results conducted within the scope of UH-60A Airloads Survey Program. The lift – propulsion plots along with the power – lift plots are compared with the wind tunnel test results for different shaft tilt angles and collective values. The comparison is made for 0.100 and 0.175 advance ratio values and both showed a satisfying correlation. Following the verification of the model, rotor maps are generated for different forward flight velocities, altitude and temperature values. The maps give information about the rotor lift and propulsion capacity since they include collective sweeps for different shaft tilt angles. Once the maps are generated for different advance ratios, oscillatory pitch link loads are probed from the maps, at the point where the helicopter weight (blade loading C_T/σ) and parasite drag (C_X/σ) intersects. The oscillatory loads at these intersection points are nondimensionalized with flight density, blade tip speed and rotor disk area to obtain C_PP. This nondimensional oscillatory load coefficient, C_PP is plotted against forward flight velocity and the divergence of the coefficient value as the forward speed increases, is observed. A linear trend for low speed is generated and the onset of stall is defined as four times of the linear trend. Another aspect to determine the onset boundary which is proposed by Lau for Lynx helicopter from its performance tests is utilized. Two methods to determine stall onset points show similar results. Finally, stall boundaries for the two methods are plotted on a blade loading (C_T/σ) vs advance ratio graph along with a comparison from literature data which includes McHugh rectangular blade stall and S70A main rotor stall. The boundary obtained with comprehensive modelling shows are good correlation with the data from literature. Looking forward, rotorcraft technology aims to achieve higher airspeeds by reducing drag and unloading the rotor by using lift-sharing wings on the rotorcrafts. Furthermore, future designs (FLRAA, S97, Bell 360 Invictus, Advanced AH-64 Compound) employ reduced drag fuselage designs together with propeller and wing to produce propulsive force and extra lift respectively. Lift-sharing wing geometries reduce the main rotor's load, C_T/σ, and improve overall aerodynamic efficiency. The fuselages with lower parasite drag shifts the rotor stall onset boundary outward. The effects of these design choices on the stall curve are thoroughly investigated, highlighting potential improvements in rotorcraft performance through aerodynamic refinements.

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